The demand for faster and more efficient aircraft is balanced with the cost of producing sufficient thrust to overcome the related drag at the desired speeds. When aircraft operate in a transonic or a supersonic regime, a flow 10 over a surface 12 of a wing 14 can form a strong normal shockwave 16 and a boundary layer 18, as shown in FIG. 1. At transonic and supersonic speeds, the strong normal shockwave 16 with its attendant wave drag being the dominant component of drag can adversely impact the aerodynamic performance of the aircraft. For example, the strong normal shockwave 16 can cause the boundary layer 18 to separate from a portion of the wing 14 and can establish a separation region 20 in the flow 10. The separation of the boundary layer 18 can further increase the drag experienced by the wing 14.
Different active or passive techniques have been investigated for control of the interaction between the strong normal shockwave 16 and the boundary layer 18. With reference to FIG. 2, one passive technique can include the use of fixed physical surface bumps 22 in the flow 10 that can tailor a local curvature of the surface 12 on the wing 14. With reference to FIG. 3, and for certain values of Mach number and angle of attack, the physical surface bumps 22 can be shown to relatively weaken the strong normal shockwave 16 and the magnitude of the associated drag. The fixed physical surface bumps 22 can establish a weak oblique shockwave 24 that can interact with the strong normal shockwave 16 to reduce its strength. At different but otherwise useful values of Mach number and angles of attack, however, the fixed physical surface bumps 22 can be shown to penalize the aerodynamic performance of the wing 14.
Other passive techniques not specifically illustrated can include sub-boundary layer vortex generators (i.e., small vertical tabs that extend into the boundary layer 18 from the surface 12 of the wing 14) and streamwise surface slits (i.e., vented wells), each of which can be positioned upstream of the strong normal shockwave 16. Each technique, at certain values of Mach number and angle of attack, can be shown to reduce the strength of the strong normal shockwave 16 and the associated wave drag. In addition, use of a porous surface (e.g., a plenum) at the foot of the strong normal shockwave 16 for certain values of Mach number and angle of attack can also be shown to reduce the strength of the strong normal shockwave and the associated wave drag by reducing a pressure jump across a foot 26 of the strong normal shockwave 16.
One active flow control technique can include a steady blowing jet located just upstream of the foot 26 of the strong normal shockwave 16 to again reduce its strength. While useful in some implementations, the steady blowing jet can be shown to reduce the lifting capability of the wing because of the relatively high jet momentum and mass flow rate required. In this instance, the Mach number of the airflow from the steady blowing jet can be comparable to the Mach number of the local airflow 10 over the wing 14. While the above examples remain useful in certain instances, there remains room in the art for improvement.